U.S. patent number 5,533,331 [Application Number 08/249,429] was granted by the patent office on 1996-07-09 for safe propulsion system for missile divert thrusters and attitude control thrusters and method for use of same.
This patent grant is currently assigned to Kaiser Marquardt, Inc.. Invention is credited to John G. Campbell, Daniel W. Ruttle.
United States Patent |
5,533,331 |
Campbell , et al. |
July 9, 1996 |
Safe propulsion system for missile divert thrusters and attitude
control thrusters and method for use of same
Abstract
A propulsion system for use with a missile or like aerial
projectile is disclosed which is suitable for use to operate the
divert thrusters and the attitude control thrusters of such a
missile while using non-toxic propellants which are entirely
non-reactive during storage, transportation, and handling. In the
preferred embodiment of the present invention, highly refined
liquid hydrocarbon fuel and oxygen gas are used as the propellants,
with a relatively small amount of the liquid hydrocarbon fuel and a
relatively large amount of the oxygen gas being combusted in an
oxygen heater to produce a hot oxygen gas containing only small
amounts of the products of combustion. The liquid hydrocarbon fuel
and the hot oxygen gas are burned in divert thrusters, and,
optionally, in attitude control thrusters to produce thrust. The
attitude control thrusters can instead alternately use either
(cold) oxygen gas from the oxygen gas storage tanks, or hot oxygen
gas from the oxygen heater to produce thrust in operation as
jets.
Inventors: |
Campbell; John G. (Northridge,
CA), Ruttle; Daniel W. (Panorama City, CA) |
Assignee: |
Kaiser Marquardt, Inc. (Foster
City, CA)
|
Family
ID: |
22943449 |
Appl.
No.: |
08/249,429 |
Filed: |
May 25, 1994 |
Current U.S.
Class: |
60/204;
60/260 |
Current CPC
Class: |
B64G
1/26 (20130101); B64G 1/401 (20130101); B64G
1/402 (20130101); F02K 9/425 (20130101) |
Current International
Class: |
B64G
1/40 (20060101); B64G 1/26 (20060101); B64G
1/22 (20060101); B64G 1/24 (20060101); F02K
9/42 (20060101); F02K 9/00 (20060101); F02K
009/42 () |
Field of
Search: |
;60/39.06,39.511,39.821,39.824,204,260,257 |
References Cited
[Referenced By]
U.S. Patent Documents
Other References
Hoyte et al. Complete Car Care Manual Reader's Digest, Inc. 1981;
p. 43. .
Armstring et al. The Diesel Engine Macmillan Co. 1959; p.
36..
|
Primary Examiner: Casaregola; Louis J.
Attorney, Agent or Firm: Posta, Jr.; John J.
Claims
What is claimed is:
1. A method of igniting fuel fluid in a propulsion system,
comprising:
a) storing a supply of fuel fluid in a fuel storage tank;
b) storing a supply of oxidant fluid in an oxidant storage
tank;
c) heating the oxidant fluid above the ignition temperature of the
fuel fluid in an oxidant heater; and
d) supplying heated oxidant fluid from said oxidant heater and said
fuel fluid from said fuel storage tank to a combustion chamber
wherein said heated oxidant fluid and said fuel fluid impinge to
thereby ignite the fuel fluid and produce thrust.
2. A method gas defined in claim 1, wherein said fuel fluid
comprises white mineral oil.
3. An ignition assembly for a propulsion system, comprising:
a) a fuel storage tank for storing a supply of fuel fluid;
b) an oxidant storage tank for storing a supply of oxidant
fluid;
c) an oxidant heater for pre-heating at least a portion of said
oxidant fluid from said oxidant storage tank to a temperature above
the ignition temperature of the fuel fluid;
d) a combustion chamber; and
e) means for supplying pre-heated oxidant fluid from said oxidant
heater and said fuel fluid from said fuel storage tank to said
combustion chamber wherein said pre-heated oxidant fluid and said
fuel fluid impinge to thereby ignite the fuel fluid and produce
thrust.
4. An ignition assembly as defined in claim 3, wherein said fuel
fluid comprises liquid hydrocarbon fuel.
5. An ignition assembly as defined in claim 4, wherein said fuel
fluid comprises white mineral oil.
6. An ignition assembly as defined in claim 3, wherein said oxidant
fluid comprises oxygen gas.
7. A propulsion system as defined in claim 3, wherein said oxidant
heater comprises:
a metal pressure vessel reinforced with a high temperature
overwrap;
means for admitting at least a portion of said oxidant fluid into
said metal pressure vessel;
means for admitting at least a portion of said fuel fluid into said
metal pressure vessel; and
means for initiating combustion within said metal pressure
vessel.
8. A propulsion system as defined in claim 7, wherein said metal
pressure vessel is made from a material with excellent oxidation
resistance at high temperatures.
9. A propulsion system as defined in claim 7, wherein said
combustion initiating means comprises:
a solid propellant initiator.
10. A propulsion system as defined in claim 7, wherein said
combustion initiating means comprises:
a spark plug.
11. A propulsion system as defined in claim 3, wherein said fuel
fluid and said oxidant fluid each comprise:
a non-toxic propellant which is non-reactive during storage,
transportation, and handling.
Description
BACKGROUND OF THE INVENTION
Field of the Invention
The present invention relates generally to safe propulsion systems
for use with a missile or like aerial projectile, and more
particularly to an improved safe propulsion system which is
suitable for use to operate the divert thrusters and the attitude
control thrusters of such a missile while using non-toxic
propellants which are entirely non-reactive during storage,
transportation, and handling.
The field of missile science has advanced rapidly during the latter
half of the twentieth century from its relatively primitive
beginnings. Early guided missiles were essentially experimental,
pilotless aircraft which were operated by radio control systems.
The tremendous technological advances in electronics have been
accompanied by similar advances in other essential fields such as
rocket propulsion, inertial guidance and control systems,
aerodynamics, material sciences, and radar systems. As a result,
guided missiles today are mass-manufactured for a variety of
purposes, ranging from military applications to carrying scientific
instruments for use in gathering information at high altitudes,
either within or above the earth's atmosphere.
While such guided missiles may vary considerably both in
application as well as size, they all include three essential
components: a power source for propelling them, a mission payload
which is to be carried by the missile, and a guidance and control
system. The first of these components is the power source, which
may be either a self-contained rocket engine, or an air-breathing
jet engine, depending on the application of the missile and
intended altitude at which the missile is intended to fly. The
second of the aforementioned components is the mission payload,
which, as mentioned above, may vary widely, varying from scientific
instruments, to surveillance equipment, to explosive warheads.
It is the third of the three essential components of a missile,
namely the guidance and control system, which is the focus of the
present invention. The internal guidance and control system of a
missile includes two elements, the first being the "brains" of the
guidance and control system, or the inertial navigation system of
the missile. The second element of the guidance and control system
is the apparatus which is used to produce the force necessary to
guide the missile in its course. While in small missiles this force
may be produced by moveable fins and other similar airfoils, in
many missiles this apparatus typically includes the divert
thrusters and the attitude control thrusters.
The divert thrusters are capable of producing a substantial amount
of thrust which is used to effect substantial course changes,
generally in two axes which are each orthogonal to the main
longitudinal axis of the missile. The attitude control system
thrusters, on the other hand, are used to effect a much finer
degree of control on the missile, rolling it around its main
longitudinal axis in either direction, or making fine course
changes in one or more directions orthogonal to the main
longitudinal axis of the missile. The divert thrusters and the
attitude control thrusters are thereby used to effect control on
the course of the missile, as directed by the inertial navigation
system of the missile.
Three different types of propulsion systems have been predominantly
utilized in the past for divert and attitude control systems used
in missiles. These three propulsion systems are liquid bipropellant
systems, liquid or gaseous monopropellant systems, and solid
propellant systems. Although these three types of systems are all
presently utilized, each of them has substantial disadvantages
which are inherent in their use, as will become evident in the
discussion to follow.
The first of the three systems widely utilized is the bipropellant
system, which uses a distinct fuel and a distinct oxidizer.
Typically, such liquid bipropellant systems use hydrazine or
monomethylhydrazine as the fuel, and nitrogen tetraoxide as the
oxidizer. In some applications, bipropellant systems use gels
instead of liquids.
Such bipropellant systems present an extreme disadvantage in that
they are highly toxic, and, as such, are completely unsatisfactory
for applications requiring non-toxicity. In addition, the
bipropellant systems are subject to detonation and rapid combustion
when dropped or exposed to a fire, making them at least potentially
highly dangerous. Finally, bipropellant systems are also dangerous
in military field applications, where a stray bullet can
potentially destroy the missile and everything in close proximity
to it.
Liquid or gaseous monopropellant systems typically also use
hydrazine as a monopropellant fuel. Such monopropellant hydrazine
is also highly toxic, and once again hydrazine is unsatisfactory
for use in applications which require non-toxicity. Monopropellant
hydrazine systems are also subject to detonation and rapid
combustion when dropped or exposed to a fire, or when hit by a
stray bullet. Hydrazine monopropellants are also subject to
detonation when heated to approximately 550.degree. F. As such,
monopropellant systems present many of the same disadvantages as
bipropellant systems.
The third type of propulsion system is the solid propellant system,
which presents an advantage over the aforementioned liquid
bipropellant and monopropellant systems in that solid propellant
systems are relatively non-toxic. Solid propellant systems will,
however, detonate when exposed to a fire or hit by a stray bullet.
In addition, solid propellant systems present several significant
disadvantages not found in liquid bipropellant and monopropellant
systems.
First, solid propellant systems are not capable of efficient on-off
pulsing operation, which presents a heavy disadvantage when solid
propellant systems are used in divert and attitude control
propulsion systems. In addition, solid propulsion systems have
other significant operational disadvantages, such as relatively
heavy weight and a poor ability to allow control of the system
center-of-gravity. As such, solid propulsion systems are even more
disadvantageous in operation than are the aforementioned liquid
bipropellant and monopropellant systems.
It is accordingly the primary objective of the present invention
that it present an improved propulsion system useable for divert
and attitude control systems, and a related method for use thereof,
which system and method use non-toxic propellants exclusively. As
such, it is a further objective of the present invention that both
the fuel and the oxidizer be non-toxic to thereby eliminate one of
the most serious drawbacks of previously known liquid or gel
bipropellants or monopropellants. It is an additional objective of
the propulsion system and the related method of the present
invention that the propellants not be subject to detonation in a
fire, when hit by a stray bullet, or when subjected to high
temperature.
It is a further objective of the propulsion system and the related
method of the present invention that the propulsion system be
efficiently operable in an on-off pulsatile manner. It is yet
another objective of the present invention that the improved
propulsion system present excellent center-of-gravity control such
that the characteristics of the missile in which the propulsion
system is installed will feature excellent dynamic
center-of-gravity characteristics. It is still another objective of
the present invention that the weight of the improved propulsion
system be relatively light in comparison to solid propulsion
systems, and comparable to or less than the weight of previously
known liquid bipropellant or monopropellant systems.
The propulsion system of the present invention must also be of a
construction which is both durable during operation, and long
lasting in a storage situation, and it should further require
little or no maintenance to be provided throughout the time that it
is stored. In order to compete effectively with previously known
liquid bipropellant and monopropellant propulsion systems, and with
solid propellant propulsion systems, the propulsion system of the
present invention should be of comparable cost to these previously
known systems, or less, to thereby afford it the broadest possible
market. Finally, it is also an objective that all of the aforesaid
advantages and objectives of the propellant system and the related
method of use of the present invention be achieved without
incurring any substantial relative disadvantage.
SUMMARY OF THE INVENTION
The disadvantages and limitations of the background art discussed
above are overcome by the present invention. With this invention, a
propulsion system and a related method for the use of the
propulsion system are presented which present a number of novel
features and advantages. First and foremost, the propulsion system
and the related method of the present invention use entirely
non-toxic propellants, the use of which to power such propulsion
systems is heretofore unknown in the art. Second, the apparatus of
the propulsion system of the present invention which allows the use
of the novel propellants, together with the propulsion system's
method of operation, are entirely novel.
The foundation of the present invention is the use of a highly
refined liquid hydrocarbon fuel and oxygen gas as the propellants.
In the preferred embodiment, the liquid hydrocarbon fuel used by
the present invention is white mineral oil, a fuel previously
unused as a propellant in missile propulsion systems. White mineral
oil is as safe a fuel as can be found; its use as baby oil and in
various food products is certainly sufficient testament to this
fact. Similarly, oxygen is an element which is necessary for life,
and as such is neither toxic nor corrosive in and of itself. The
oxygen gas may be stored in oxygen storage tanks at high pressure,
typically approximately 10,000 psia, while the liquid hydrocarbon
fuel is stored in fuel storage tanks at low pressure, typically
approximately one atmosphere.
It will at once be appreciated by those skilled in the art that the
use of the novel propellants of the present invention presents
significant advantages in storage, since the propellants of the
present invention are neither as volatile nor as toxic as
previously known propellants. However, due to the nature of the
novel propellants utilized by the present invention, a novel
propulsion system structure is certainly necessary to bring about
the successful use of the propellants of the present invention. As
this structure is discussed herein, it will also be appreciated by
those skilled in the art that while the propulsion system of the
present invention is remarkably novel, it is certainly neither
difficult nor expensive to implement.
If the novel combination of propellants discussed above is the
foundation upon which the present invention builds, then the heart
of the present invention is an oxygen heater, which is used to
produce heated oxygen gas at a temperature of approximately
1200.degree. F. to 2000.degree. F. The oxygen heater is supplied
with oxygen gas at a pressure of approximately 1000 psi, and a
small amount of the liquid hydrocarbon fuel at a similar pressure.
Combustion in the oxygen heater may be initiated by a solid
propellant initiator (which is the implementation chosen in the
preferred embodiment), or, alternately, by a spark plug which may
be used in addition to or instead of the solid propellant
initiator.
Since a relatively large amount of oxygen gas is introduced into
the oxygen heater with respect to the relatively small amount of
the liquid hydrocarbon fuel introduced into the oxygen heater
(typically approximately a 40:1 mass flow ratio), the gas stream
leaving the oxygen heater contains only small amounts of products
of combustion. Thus, the primary gaseous product leaving the oxygen
heater will be oxygen gas (approximately 89 percent of the total
gaseous products leaving the oxygen heater). The gas temperature in
the oxygen heater for the 40:1 mass flow ratio will be
approximately 1850.degree. F., which temperature will be maintained
by the oxygen heater even with the cool oxygen gas and the cool
liquid hydrocarbon fuel entering the oxygen heater.
The liquid hydrocarbon fuel may be pressurized by using gas
pressure acting on a piston contained in each of the fuel storage
tanks containing the liquid hydrocarbon fuel. Pressurization of the
fuel storage tanks may be generated either by an independent gas
supply such as from high pressure helium tanks, or by cold oxygen
gas supplied from the oxygen storage tanks. In the latter case, a
burst disk is used to isolate the oxygen gas from the fuel storage
tanks during storage, and this burst disk will remain intact until
just before the propulsion system is to be used.
The propulsion system of the present invention may advantageously
be used to operate the divert thrusters and the attitude control
thrusters of a missile. The divert thrusters are bipropellant
thrusters, with the propellants being the liquid hydrocarbon fuel
together with the hot oxygen gas from the oxygen heater. Each of
the divert thrusters has a pair of control valves, to respectively
control the supply of liquid hydrocarbon fuel and the supply of hot
oxygen gas to that particular divert thruster.
Ignition of the propellants in the divert thrusters of the present
invention is accomplished by injecting the liquid hydrocarbon fuel
into the hot oxygen gas stream. In a first embodiment, a single
stage divert thruster is taught in which both ignition and
combustion occur in a single ignition/combustion chamber. In a
second embodiment, a two stage divert thruster is taught in which
ignition occurs in an ignition chamber, with the primary combustion
taking place in an adjacent downstream combustion chamber.
The attitude control thrusters of the preferred embodiment (which
in all but one of the embodiments herein are properly referred to
as jets rather than monopropellant thrusters) use only pressurized
oxygen gas to produce thrust. In a first series of embodiments, the
attitude control thrusters use cold oxygen gas from the oxygen
storage tanks. In a second series of embodiments, the attitude
control thrusters use hot oxygen gas from the oxygen heater. In
each of these embodiments, the flow of oxygen to each attitude
control thruster is controlled by a valve.
In an alternate embodiment, the attitude control thrusters are
bipropellant thrusters, with the propellants being the liquid
hydrocarbon fuel and the hot oxygen gas. In this embodiment, each
of the attitude control thrusters has a pair of control valves, to
respectively control the supply of liquid hydrocarbon fuel and the
supply of hot oxygen gas to that particular attitude control
thruster. Ignition of the propellants in the attitude control
thrusters of this embodiment is accomplished by injecting the
liquid hydrocarbon fuel into the hot oxygen gas stream in a manner
similar to that described above with regard to the divert
thrusters.
It may therefore be seen that the present invention teaches an
improved propulsion system useable for divert and attitude control
propulsion systems, and a related method for use thereof, which
system and method exclusively use propellants which are non-toxic.
As such, both the fuel and the oxidizer of the present invention
are non-toxic, thereby eliminating one of the most serious
drawbacks of previously known liquid or gel bipropellants or
monopropellants. The propellants used by the propulsion system and
the related method of the present invention are not subject to
detonation in a fire, when hit by a stray bullet, or when subjected
to high temperature.
The propulsion system and the related method of the present
invention are efficiently operable in an on-off pulsatile manner,
unlike solid propellant systems. The improved propulsion system of
the present invention also presents excellent center-of-gravity
control, such that the characteristics of the missile in which the
propulsion system of the present invention is installed will also
feature excellent dynamic center-of-gravity characteristics. The
weight of the improved propulsion system of the present invention
is also relatively light in comparison to solid propulsion systems,
and in fact is comparable to or less than the weight of previously
known liquid bipropellant or monopropellant systems.
The propulsion system of the present invention is also of a
construction which is both durable during operation, and long
lasting in a storage situation, and which requires little or no
maintenance to be provided throughout the time that it is stored.
The propulsion system of the present invention is of comparable
cost to previously known liquid bipropellant and monopropellant
propulsion systems, and to solid propellant propulsion systems,
thereby affording the propulsion system of the present invention
the broadest possible market. Finally, all of the aforesaid
advantages and objectives of the propellant system and the related
method of use of the present invention are achieved without
incurring any substantial relative disadvantage.
DESCRIPTION OF THE DRAWINGS
These and other advantages of the present invention are best
understood with reference to the drawings, in which:
FIG. 1 is a somewhat schematic cross-sectional view of a missile,
showing a pair of liquid hydrocarbon fuel storage tanks, a pair of
oxygen gas storage tanks, an oxygen heater, four divert thrusters
respectively located at the left side, the top, the bottom, and the
right side of the missile and oriented radially outwardly, and six
attitude control thrusters with two opposing pairs of the attitude
control thrusters being located at the left side and the right side
of the missile and oriented upwardly and downwardly, and with
single attitude control thrusters being located at the top and the
bottom of the missile and oriented radially outwardly;
FIG. 2 is a functional schematic of the oxygen heater illustrated
in FIG. 1, showing a solid propellant initiator and an optional
spark plug located inside a metal pressure vessel and its high
temperature overwrap, and also showing a projectile control system
operating the solid propellant initiator and the spark plug, and
additionally showing the supply of liquid hydrocarbon fuel and the
supply of oxygen gas to the oxygen heater, and the supply of liquid
hydrocarbon fuel and the supply of hot oxygen gas which will be
provided to divert thrusters;
FIG. 3 is a somewhat schematic, partially cutaway view of a single
stage divert thruster having a single stage in which both ignition
and combustion occur, and showing the supply valves through which
liquid hydrocarbon fuel and hot oxygen gas are supplied to the
divert thruster;
FIG. 4 is a somewhat schematic, partially cutaway view of a two
stage divert thruster having a first stage in which ignition
occurs, and a second stage in which combustion occurs, and showing
the supply valves through which liquid hydrocarbon fuel and hot
oxygen gas are supplied to the divert thruster;
FIG. 5 is a functional schematic view of a first propulsion system
constructed according to the teachings of the present invention,
with oxygen gas from an oxygen plenum being used to pressurize the
liquid hydrocarbon fuel storage tanks, and with oxygen gas from the
oxygen plenum being supplied through valves to six attitude control
thrusters;
FIG. 6 is a functional schematic view of a second propulsion system
constructed according to the teachings of the present invention,
with oxygen gas from an oxygen plenum being used to pressurize the
liquid hydrocarbon fuel storage tanks, and with hot oxygen gas from
the oxygen heater being supplied through valves to six attitude
control thrusters;
FIG. 7 is a functional schematic view of a third propulsion system
constructed according to the teachings of the present invention,
with oxygen gas from an oxygen plenum being used to pressurize the
liquid hydrocarbon fuel storage tanks, and with hot oxygen gas from
the oxygen plenum and liquid hydrocarbon fuel being supplied
through valves to six attitude control thrusters;
FIG. 8 is a functional schematic view of a fourth propulsion system
constructed according to the teachings of the present invention,
with helium gas from helium gas storage tanks being used to
pressurize the liquid hydrocarbon fuel storage tanks, and with
oxygen gas from an attitude control system plenum being supplied
through valves to six attitude control thrusters;
FIG. 9 is a functional schematic view of a fifth propulsion system
constructed according to the teachings of the present invention,
with helium gas from helium gas storage tanks being used to
pressurize the liquid hydrocarbon fuel storage tanks, and with
oxygen gas from the oxygen plenum being supplied through valves to
six attitude control thrusters; and
FIG. 10 is a functional schematic view of a sixth propulsion system
constructed according to the teachings of the present invention,
with helium gas from helium gas storage tanks being used to
pressurize the liquid hydrocarbon fuel storage tanks, and with hot
oxygen gas from the oxygen heater being supplied through valves to
six attitude control thrusters.
DETAILED DESCRIPTION OF THE PREFERRED EMBODIMENT
The preferred embodiment differs from previously known propulsion
systems and methods of use in two key respects. First, the
propellants used by the propulsion system and the related method of
use of the present invention differ greatly from previously known
propellants; the preferred propellants of the present invention are
highly refined liquid hydrocarbon fuel, such as, for example, white
mineral oil, and oxygen gas. Secondly, the propulsion system and
the related method of use of the present invention use an oxygen
heater to heat the oxygen gas prior to providing it to thrusters
where it is mixed with the liquid hydrocarbon fuel and burned.
These two key aspects of the present invention provide significant
advantages in their use in a propulsion system. While such a
propulsion system has various applications, the example used herein
describes the use of the propulsion system and the related method
of use of the present invention in an application including both
divert thrusters and attitude control thrusters. Other varied
applications of the apparatus and the method of the present
invention will be apparent to those skilled in the art after
reviewing the following exemplary description.
Referring first to FIG. 1, the principal components of the present
invention are schematically illustrated in position within the
cross-section of a missile 30. The missile 30 is schematically
illustrated as a structural tube 32, which is typically made of a
high strength, light weight material such as graphite epoxy.
Located within the structural tube 32 in the missile 30 is a
structural bulkhead 34, on the surface of which structural bulkhead
34 the various principal components of the present invention are
mounted. It will be understood by those skilled in the art that
these components are mounted within the structural tube 32 of the
missile 30.
Centrally located in the four quadrants of the missile 30 are four
propellant storage tanks. Located in the upper left and lower right
quadrants (as illustrated in FIG. 1) are two liquid hydrocarbon
fuel storage tanks 36 and 38, respectively. The liquid hydrocarbon
fuel storage tanks 36 and 38 are designed to hold liquid
hydrocarbon fuel at atmospheric pressure.
The liquid hydrocarbon fuel storage tanks 36 and 38 are designed to
be connected in parallel to supply the propulsion system of the
present invention with liquid hydrocarbon fuel in order to
establish and maintain a mass balance about the center of gravity.
(As will become evident below in conjunction with the discussion of
FIGS. 5 through 10, the liquid hydrocarbon fuel storage tanks 36
and 38 also contain a piston mechanism designed to keep the liquid
hydrocarbon fuel contained therein at one end of the liquid
hydrocarbon fuel storage tanks 36 and 38.)
Located in the upper right and lower left quadrants (as illustrated
in FIG. 1) are two oxygen gas storage tanks 40 and 42,
respectively. The oxygen gas storage tanks 40 and 42 are designed
to hold oxygen gas at a very high pressure, such as, for example,
approximately 10,000 psia. The oxygen gas storage tanks 40 and 42
may be made of or be reinforced by a high strength, light weight
material such as graphite epoxy. The oxygen gas storage tanks 40
and 42 are also designed to be connected in parallel to supply the
propulsion system of the present invention with oxygen gas in order
to establish and maintain a mass balance about the center of
gravity.
Located in the center of the structural bulkhead 34 in the missile
30 as illustrated in FIG. 1 is an oxygen heater 44. The oxygen
heater 44 will function as a gas generator to generate hot oxygen
gas, which will be used together with the liquid hydrocarbon fuel
as propellants for the divert thrusters of the propulsion system of
the present invention. In several of the embodiments of the present
invention to be discussed below, the hot oxygen gas generated by
the oxygen heater 44 will also be used by the attitude control
thrusters of the propulsion system of the present invention to
generate thrust.
Located at ninety degree intervals spaced around the missile 30 on
the structural bulkhead 34 of the missile 30 are four divert
thrusters 46, 48, 50, and 52. The divert thruster 46 is located at
the left side of the missile 30 (as illustrated in FIG. 1), the
divert thruster 48 is located at the top of the missile 30 (as
illustrated in FIG. 1), the divert thruster 50 is located at the
bottom of the missile 30 (as illustrated in FIG. 1), and the divert
thruster 52 is located at the right side of the missile 30 (as
illustrated in FIG. 1). The four divert thrusters 46, 48, 50, and
52 are each oriented radially outwardly with respect to the
longitudinal axis of the missile 30 (which runs through the center
of the oxygen heater 44).
Six attitude control thrusters 54, 56, 58, 60, 62, and 64 are also
located on the structural bulkhead 34 of the missile 30. The
attitude control thrusters 54 and 56 are located at the left side
of the missile 30 (as illustrated in FIG. 1), with the attitude
control thruster 54 being oriented upwardly, and the attitude
control thruster 56 being oriented downwardly. The attitude control
thruster 58 is located at the top of the missile 30 (as illustrated
in FIG. 1) and is oriented upwardly, and the attitude control
thruster 60 is located at the bottom of the missile 30 (as
illustrated in FIG. 1) and is oriented downwardly. The attitude
control thrusters 62 and 64 are located at the right side of the
missile 30 (as illustrated in FIG. 1), with the attitude control
thruster 62 being oriented upwardly, and the attitude control
thruster 64 being oriented downwardly.
Referring next to FIG. 2, the oxygen heater 44 of FIG. 1 is
illustrated in a functional schematic manner which illustrates its
interaction with the rest of the propulsion system of the present
invention. The operation of the propulsion system is controlled by
a projectile computer control system 66, which is also shown
schematically in FIG. 2.
The oxygen heater 44 consists essentially of a metal pressure
vessel 68 which is wrapped with a high temperature overwrap 70. The
metal pressure vessel 68 is typically made of a material which has
excellent oxidation resistance at high temperature, such as, for
example, a nickel alloy, a cobalt alloy, or iridium. The metal
pressure vessel 68 must be capable of easily withstanding the high
operational temperature of the oxygen heater 44, which is
approximately 1850.degree. F.
Oxygen gas from the oxygen gas storage tanks 40 and 42 (FIG. 1) and
liquid hydrocarbon fuel from the liquid hydrocarbon fuel storage
tanks 36 and 38 (FIG. 1) are supplied under pressure (the details
of which will be discussed later in conjunction with FIG. 5) to the
oxygen heater 44. Injection of the oxygen gas is distributed
throughout the volume of the interior of the oxygen heater 44 in
the preferred embodiment to minimize the quenching effect of the
cold oxygen gas entering the oxygen heater 44.
The oxygen heater 44 of the preferred embodiment uses a solid
propellant initiator 72 to generate hot gas at the beginning of the
operation of the oxygen heater 44. If the solid propellant
initiator 72 is utilized in the oxygen heater 44, it is capable of
generating hot gas from the oxygen heater 44 even prior to the
pressurized supply of oxygen gas and liquid hydrocarbon fuel being
supplied to the oxygen heater 44.
Ignition of the solid propellant initiator 72 is initiated by the
projectile computer control system 66. It will become apparent
below in conjunction with the discussion of FIG. 5 that the
projectile computer control system 66 also controls the pressurized
supply of oxygen gas and liquid hydrocarbon fuel to the oxygen
heater 44.
Also schematically shown in FIG. 2 is a spark plug 74, which may
optionally be used to initiate or maintain ignition of oxygen gas
and liquid hydrocarbon fuel within the oxygen heater 44 by spark
ignition. The spark plug 74 is typically used to maintain ignition
of oxygen gas and liquid hydrocarbon fuel within the oxygen heater
44, and can only initiate ignition when pressurized oxygen gas and
liquid hydrocarbon fuel are already being supplied to the oxygen
heater 44.
In any event, following ignition, the oxygen heater 44 will produce
a pressurized flow of hot oxygen gas at a temperature of
approximately 1200.degree. F. to 2000.degree. F. Despite the fact
that combustion takes place within the oxygen heater 44, only small
amounts of the products of combustion are contained in the hot gas
generated by the oxygen heater 44. For example, assuming an oxygen
to fuel mass flow ratio of 40:1 into the oxygen heater 44, the hot
gas produced by the oxygen heater 44 will consist of approximately
89 percent oxygen (in the form of O.sub.2), 5.5 percent water
H.sub.2 O), and 5.5 percent carbon dioxide (CO.sub.2). For this
oxygen to fuel mass flow ratio, the gas temperature within the
oxygen heater 44 will be approximately 1850.degree. F.
Since the gas stream generated by the oxygen heater 44 is
predominantly oxygen, it is referred to in this specification as
hot Oxygen gas even though it does contain small amounts of the
other materials mentioned above. Thus, the hot oxygen gas and the
liquid hydrocarbon fuel are supplied to the divert thrusters 46,
48, 50, and 52 (illustrated in FIG. 1).
In accordance with the preferred embodiment of the propulsion
system and the method of the present invention, two divert thruster
designs are presented herein. The first of these divert thrusters
is a single stage divert thruster having a single stage in which
both ignition and combustion occur, while the second of these
divert thrusters is a two stage divert thruster having a first
stage in which ignition occurs, and a second stage in which
combustion primarily occurs.
Referring now to FIG. 3, the first of these divert thruster designs
is presented. A single stage divert thruster 76 in which both
ignition and combustion occur in a single stage is illustrated in
FIG. 3. The single stage divert thruster 76 consists of a housing
member 78, which is shown as a single segment in the simplified
drawing of FIG. 3. In reality, the housing member 78 would include
a number of different components, the general construction of which
is readily apparent to those skilled in the art.
Located inside the housing member 78 at the left side thereof as
shown in FIG. 3 is an oxygen gas inlet chamber 80. Hot oxygen gas
is supplied to the oxygen gas inlet chamber 80 through an oxidizer
valve 82 whenever the oxidizer valve 82 is opened. When hot oxygen
gas is allowed to enter the oxygen gas inlet chamber 80 through the
oxidizer valve 82, it can only exit the oxygen gas inlet chamber 80
through an oxygen gas injector orifice 84, from which it enters at
the left end of a cylindrical ignition/combustion chamber 86 which
is located inside the housing member 78. It will be appreciated by
those skilled in the art that when the oxidizer valve 82 is opened,
hot oxygen gas will flow through the oxygen gas injector orifice 84
into the ignition/combustion chamber 86 at high speed.
Also located inside the housing member 78 at the left side and near
the top thereof as shown in FIG. 3 is a small fuel inlet chamber
88. Liquid hydrocarbon fuel is supplied to the fuel inlet chamber
88 through a fuel valve 90 whenever the fuel valve 90 is opened.
The right side of the fuel inlet chamber 88 is in fluid
communication with an annular fuel chamber 92, which surrounds and
is spaced away from the oxygen gas injector orifice 84.
Some of the liquid hydrocarbon fuel in the annular fuel chamber 92
is injected through fuel orifices 94 into the flow of hot oxygen
gas flowing through the oxygen gas injector orifice 84. The liquid
hydrocarbon fuel so injected into the hot oxygen gas flow in the
oxygen gas injector orifice 84 will create a mixture of small fuel
droplets and oxygen. The fuel orifices 94 may either be single or
doublet orifices.
Some of the liquid hydrocarbon fuel contained in the annular fuel
chamber 92 will be injected through orifices 96 leading into the
ignition/combustion chamber 86, at locations which are not directly
in the flow path of the hot oxygen gas from the oxygen gas injector
orifice 84. The liquid hydrocarbon fuel injected through the
orifices 96 will initially act to cool the interior walls of the
housing member 78 which form the ignition/combustion chamber 86,
and will then form droplets passing into the ignition/combustion
chamber 86 itself.
As the liquid hydrocarbon fuel is injected into the hot oxygen gas
flow, it will be ignited by contact with the hot oxygen gas itself.
It may be noted that the temperature of the hot oxygen gas is
selected to ensure ignition within the ignition/combustion chamber
86. This ignition will occur at or near the left side of the
ignition/combustion chamber 86, and the droplets of liquid
hydrocarbon fuel will be burned in the ignition/combustion chamber
86. As the gasses produced by the combustion process expand, they
pass through a throat 98 at the right end of the
ignition/combustion chamber 86 into an exhaust cone 100, from which
they exit the single stage divert thruster 76 while producing
thrust.
Referring next to FIG. 4, the second of the divert thruster designs
of the present invention is presented. A two stage divert thruster
102 in which ignition occurs in a first stage and combustion occurs
primarily in a second stage is illustrated in FIG. 4. The two stage
divert thruster 102 consists of a housing member 104, which is
shown as a single segment in the simplified drawing of FIG. 4. In
reality, the housing member 104 would include a number of different
components, the general construction of which is readily apparent
to those skilled in the art.
Located inside the housing member 104 at the left side thereof as
shown in FIG. 4 is an oxygen gas inlet chamber 106. Hot oxygen gas
is supplied to the oxygen gas inlet chamber 106 through an oxidizer
valve 108 whenever the oxidizer valve 108 is opened. When hot
oxygen gas is allowed to enter the oxygen gas inlet chamber 106
through the oxidizer valve 108, it can only exit the oxygen gas
inlet chamber 106 through an oxygen gas injector orifice 110, from
which it enters at the left end of an essentially cylindrical
ignition chamber 112 which is located inside the housing member
104. It will be appreciated by those skilled in the art that when
the oxidizer valve 108 is opened, hot oxygen gas will flow through
the oxygen gas injector orifice 110 into the ignition chamber 112
at high speed.
Also located inside the housing member 104 at the left side and
near the top thereof as shown in FIG. 4 is a small fuel inlet
chamber 114. Liquid hydrocarbon fuel is supplied to the fuel inlet
chamber 114 through a fuel valve 116 whenever the fuel valve 116 is
opened. The right side of the fuel inlet chamber 114 is in fluid
communication with an annular fuel chamber 118, which surrounds and
is spaced away from the oxygen gas injector orifice 110.
The liquid hydrocarbon fuel contained in the annular fuel chamber
118 is injected through fuel orifices 120 into the flow of hot
oxygen gas flowing through the oxygen gas injector orifice 110. The
liquid hydrocarbon fuel so injected into the hot oxygen gas flow in
the oxygen gas injector orifice 110 will create a mixture of small
fuel droplets and oxygen. The fuel orifices 120 may either be
single or doublet orifices.
As the liquid hydrocarbon fuel is injected into the hot oxygen gas
flow in the oxygen gas injector orifice 110, it will be ignited by
contact with the hot oxygen gas itself. It may be noted that the
length of the cylindrical ignition chamber 112 and the temperature
of the hot oxygen gas is selected to ensure ignition within the
cylindrical ignition chamber 112. The ignition will occur at or
near the left side of the cylindrical ignition chamber 112, and the
droplets of liquid hydrocarbon fuel will thus be ignited in the
cylindrical ignition chamber 112.
The hot gas in the cylindrical ignition chamber 112 will exit the
cylindrical ignition chamber 112 at the right end thereof, from
which it enters a larger diameter combustion chamber 122. The
mixture ratio of liquid hydrocarbon fuel and hot oxygen gas
injected into the cylindrical ignition chamber 112 is selected to
produce a sufficiently hot gas temperature which will ensure
ignition of the remainder of the liquid hydrocarbon fuel in the
combustion chamber 122, as will become evident in the following
description.
Located to the left of the left side of the combustion chamber 122
and disposed within the housing member 104 is an annular fuel
chamber 124 which surrounds and is spaced away from the right end
of the cylindrical ignition chamber 112. Liquid hydrocarbon fuel is
supplied through a supply tube 126 to the annular fuel chamber 124
whenever the fuel valve 116 is opened.
Some of the liquid hydrocarbon fuel in the annular fuel chamber 124
is injected through fuel orifices 128 into the flow of hot gas
flowing from the cylindrical ignition chamber 112 into the
combustion chamber 122. The liquid hydrocarbon fuel so injected
into the hot gas flow in the combustion chamber 122 will create a
mixture of small fuel droplets and oxygen. The fuel orifices 128
may either be single or doublet orifices.
Some of the liquid hydrocarbon fuel contained in the annular fuel
chamber 124 will be injected through orifices 130 leading into the
combustion chamber 122, at locations which are not directly in the
flow path of the hot gas entering the combustion chamber 122 from
the cylindrical ignition chamber 112. The liquid hydrocarbon fuel
injected through the orifices 130 will initially act to cool the
interior walls of the housing member 104 which form the combustion
chamber 122, and will then form droplets passing into the
combustion chamber 122 itself.
As the liquid hydrocarbon fuel is injected into the hot gas flow,
it will be ignited by contact with the hot oxygen gas itself. As
previously mentioned, the mixture ratio of liquid hydrocarbon fuel
and hot oxygen gas injected into the cylindrical ignition chamber
112 produces a sufficiently hot gas temperature to ensure ignition
of the liquid hydrocarbon fuel in the combustion chamber. This
ignition will occur at or near the left side of the combustion
chamber 122, and the droplets of liquid hydrocarbon fuel will be
burned in the combustion chamber 122. As the gasses produced by the
combustion process expand, they pass through a throat 132 at the
right end of the combustion chamber 122 into an exhaust cone 134,
from which they exit the two stage divert thruster 102 while
producing thrust.
Referring next to FIG. 5, the first of six exemplary embodiments of
the propulsion system of the present invention is illustrated.
These six exemplary embodiments each contain different
implementations presenting different combinations of features and
advantages. All six of the embodiments illustrated in FIGS. 5
through 10 contain the principal components illustrated in FIG. 1:
the liquid hydrocarbon fuel storage tanks 36 and 38, the oxygen gas
storage tanks 40 and 42, the oxygen heater 44, the divert thrusters
46, 48, 50, and 52, the attitude control thrusters 54, 56, 58, 60,
62, and 64, the projectile computer control system 66.
Referring now specifically to FIG. 5, it may be seen that the
liquid hydrocarbon fuel storage tanks 36 and 38 function as
cylinders in which pistons are located to force liquid hydrocarbon
fuel from outlets in the liquid hydrocarbon fuel storage tanks 36
and 38. Specifically, the liquid hydrocarbon fuel storage tank 36
has a piston 136 contained therein, and the liquid hydrocarbon fuel
storage tank 38 has a piston 138 contained therein. The pistons 136
and 138 in the liquid hydrocarbon fuel storage tanks 36 and 38,
respectively, are driven by pressurized gas entering through inlets
in the liquid hydrocarbon fuel storage tanks 36 and 38.
The inlets of the liquid hydrocarbon fuel storage tanks 36 and 38
are tied together, and the outlets of the liquid hydrocarbon fuel
storage tank 36 and 38 are tied together. (Note that the systems of
FIGS. 6 through 10, which will be discussed below, each have the
same interconnections and operation of their respective liquid
hydrocarbon fuel storage tanks 36 and 38.)
Referring again to FIG. 5, the outlets of the oxygen gas storage
tank 40 and 42 are tied together. (Note that the propulsion systems
of FIGS. 6 through 10, which will be discussed below, each have the
same interconnections of their respective oxygen gas storage tank
40 and 42.)
Referring once again to FIG. 5, the divert thrusters 46, 48, 50,
and 52 each have a fuel valve and an oxidizer valve. Specifically,
the divert thruster 46 has a fuel valve 140 and an oxidizer valve
148, the divert thruster 48 has a fuel valve 142 and an oxidizer
valve 150, the divert thruster 50 has a fuel valve 144 and an
oxidizer valve 152, and the divert thruster 52 has a fuel valve 146
and an oxidizer valve 154. Operation of the fuel valves 140, 142,
144, and 146 and the oxidizer valves 148, 150, 152, and 154 are all
controlled by the projectile computer control system 66. (Note that
the propulsion systems of FIGS. 6 through 10, which will be
discussed below, each have the same fuel valve and oxidizer valve
arrangement on their respective divert thrusters 46, 48, 50, and
52.)
Referring still again to FIG. 5, the attitude control thrusters 54,
56, 58, 60, 62, and 64 are jets in their operation, using only a
single pressurized fluid to provide thrust. As such, they each
require only a single control valve. Specifically, the attitude
control thruster 54 has a control valve 156, the attitude control
thruster 56 has a control valve 158, the attitude control thruster
58 has a control valve 160, the attitude control thruster 60 has a
control valve 162, the attitude control thruster 62 has a control
valve 164, and the attitude control thruster 64 has a control valve
166. Operation of the control valves 156, 158, 160, 162, 164, and
166 are all controlled by the projectile computer control system
66. (Note that the propulsion systems of FIGS. 6 and 8 through 10,
which will be discussed below, each have the same control valve
arrangement on their respective attitude control thrusters 54, 56,
58, 60, 62, and 4. Note, however, that the propulsion system of
FIG. 7 has a different attitude control thruster arrangement, which
will be specifically described below in conjunction with the
discussion of FIG. 7.)
Referring now once more to FIG. 5, as noted above the outlets of
the oxygen gas storage tanks 40 and 42 are tied together. Oxygen
gas is stored in the oxygen gas storage tanks 40 and 42 at high
pressure, typically on the order of approximately 10,000 psia.
Pressure in the oxygen gas storage tanks 40 and 42 is monitored by
a pressure transducer 168, which supplies a pressure signal which
is monitored by the projectile computer control system 66.
For safety purposes, a burst disk 170 is installed on the outlets
from the oxygen gas storage tanks 40 and 42. The burst disk 170 is
designed to rupture at a pressure higher than the maximum oxygen
gas pressure in storage or operation of the propulsion system, but
lower than the burst pressure of the oxygen gas storage tanks 40
and 42. In the event of the propulsion system being engulfed in a
fire, the pressure in the oxygen gas storage tanks 40 and 42 will
rise until the burst disk 170 ruptures. The oxygen gas will then be
discharged from the oxygen gas storage tanks 40 and 42, perhaps
temporarily increasing the intensity of the fire, but avoiding the
rupture and fragmentation of the oxygen gas storage tanks 40 and
42.
A pyrovalve 172 is installed between the outlets from the oxygen
gas storage tanks 40 and 42 and a solenoid pressure control valve
174. The pyrovalve 172 is made for one-time operation, and once
opened by a signal from the projectile computer control system 66
will remain open to allow the passage of oxygen gas therethrough.
The pyrovalve 172 functions to isolate the rest of the system from
the oxygen gas during storage.
Once the pyrovalve 172 has been opened, the solenoid pressure
control valve 174 will be opened by the projectile computer control
system 66 to allow oxygen gas to flow into an oxygen plenum 176. As
an added safety feature, note that if the pyrovalve 172 were fired
accidentally, or by the occurrence of a fire, high pressure oxygen
gas would still be isolated from the oxygen plenum 176 by the
solenoid pressure control valve 174. Pressure in the oxygen plenum
176 is monitored by a pressure transducer 178, which supplies a
pressure signal which is monitored by the projectile computer
control system 66. Using this pressure signal, the projectile
computer control system 66 will operate the solenoid pressure
control valve 174 to control the pressure of oxygen gas in the
oxygen plenum 176.
Typically, pressure in the oxygen plenum 176 is maintained at
approximately 1000 psi. Whenever the pressure in the oxygen plenum
176 drops below the desired pressure, the projectile computer
control system 66 will cause the solenoid pressure control valve
174 to be actuated for a short duration, typically a few
milliseconds, to allow more oxygen gas to be injected into the
oxygen plenum 176, thereby raising the pressure in the oxygen
plenum 176.
(Note that the propulsion systems of FIGS. 6 through 10, which will
be discussed below, each have the same arrangements of the pressure
transducer 168, the burst disk 170, the pyrovalve 172, the solenoid
pressure control valve 174, the oxygen plenum 176, and the pressure
transducer the pressure transducer 178 as contained in the
propulsion system of FIG. 5, and are operated in the same manner by
their respective projectile computer control systems 66.)
Referring once again to FIG. 5, as noted above, the inlets and the
outlets of the liquid hydrocarbon fuel storage tanks 36 and 38 are
tied together. The liquid hydrocarbon fuel, which is white mineral
oil in the preferred embodiment, is stored in the liquid
hydrocarbon fuel storage tanks 36 and 38 at low pressure, typically
on the order of approximately 1 atmosphere.
An outlet from the oxygen plenum 176 is connected to the inlets of
the liquid hydrocarbon fuel storage tanks 36 and 38 with a burst
disk 180 installed therein. The burst disk 180 functions to isolate
the oxygen plenum 176 from the liquid hydrocarbon fuel storage
tanks 36 and 38 during long term storage. The burst disk 180 is
designed to rupture at a pressure lower than the operating pressure
of the oxygen plenum 176, so that when the oxygen plenum 176 is
pressurized by initially opening the solenoid pressure control
valve 174, the burst disk 180 will rupture and allow oxygen gas to
enter the inlets of the liquid hydrocarbon fuel storage tanks 36
and 38, thereby pressurizing the liquid hydrocarbon fuel storage
tanks 36 and 38 to approximately 1000 psi.
(Note that the propulsion systems of FIGS. 6 and 7, which will be
discussed below, each have the same burst disk 180 arrangement, and
use oxygen gas from the oxygen plenum 176 to pressurize their
respective liquid hydrocarbon fuel storage tanks 36 and 38 in the
same manner as the propulsion system of FIG. 5. Note, however, that
the propulsion system of FIGS. 8 through 10 have a different
arrangement used to pressurize their respective liquid hydrocarbon
fuel storage tanks 36 and 38, which arrangement will be described
below in conjunction with the discussion of FIG. 8.)
Referring again now to FIG. 5, for safety purposes, a vent burst
disk 182 is installed on the outlets from the liquid hydrocarbon
fuel storage tanks 36 and 38. The vent burst disk 182 is designed
to rupture at a pressure higher than the gas pressure used to
pressurize the inlets of the liquid hydrocarbon fuel storage tanks
36 and 38, but lower than the burst pressure of the liquid
hydrocarbon fuel storage tanks 36 and 38. In the event of the
propulsion system being engulfed in a fire, the pressure in the
liquid hydrocarbon fuel storage tanks 36 and 38 will rise until the
vent burst disk 182 ruptures. The liquid hydrocarbon fuel will then
be discharged from the liquid hydrocarbon fuel storage tanks 36 and
38, thereby avoiding the rupture and fragmentation of the liquid
hydrocarbon fuel storage tanks 36 and 38.
When the liquid hydrocarbon fuel storage tanks 36 and 38 are
initially pressurized, a burst disk 184 will rupture, allowing
liquid hydrocarbon fuel to flow therethrough to a fuel manifold
186. Pressure in the fuel manifold 186 is thus maintained at
approximately the same level as the pressure in the oxygen plenum
176, or approximately 1000 psi. Prior to pressurization of the
liquid hydrocarbon fuel storage tanks 36 and 38, the liquid
hydrocarbon fuel is isolated in the liquid hydrocarbon fuel storage
tanks 36 and 38 by the burst disk 184. The burst disk 184 will
rupture at a pressure less than the pressure at which the liquid
hydrocarbon fuel storage tanks 36 and 38 are pressurized.
Typically, the burst disk 184 will rupture at a pressure on the
order of approximately 200 psi. (Note that the propulsion systems
of FIGS. 6 through 10, which will be discussed below, each have the
same arrangements and operational characteristics of the vent burst
disk 182 and the burst disk 184 as contained in the propulsion
system of FIG. 5.)
Referring once again to FIG. 5, the liquid hydrocarbon fuel is also
supplied under pressure from the fuel manifold 186 to a check valve
188, through which the liquid hydrocarbon fuel is injected into the
oxygen heater 44. The check valve 188 functions to ensure one-way
flow therethrough. Liquid hydrocarbon fuel is also supplied under
pressure from the fuel manifold 186 to a solenoid pressure control
valve 190, which, when actuated by the projectile computer control
system 66, will allow additional liquid hydrocarbon fuel to be
supplied through a check valve 192, through which the additional
liquid hydrocarbon fuel is injected into the oxygen heater 44.
The check valve 192 also functions to ensure one-way flow
therethrough. The details of actuation of the solenoid pressure
control valve 190 will be discussed below. (Note that the
propulsion systems of FIGS. 6 through 10, which will be discussed
below, each have the same arrangements and operational
characteristics of the check valve 188, the solenoid pressure
control valve 190, and the check valve 192 as contained in FIG.
5.)
Referring once again to FIG. 5, oxygen gas is supplied under
pressure from the oxygen plenum 176 to a check valve 194, through
which the oxygen gas is injected into the oxygen heater 44. The
check valve 194 functions to ensure one-way flow therethrough.
In the preferred embodiment, the cracking pressure differential
required to open the check valves 188 and 194 can be adjusted so
that the check valve 188 will open to allow liquid hydrocarbon fuel
into the oxygen heater 44 before the check valve 192 will open to
allow oxygen gas into the oxygen heater 44, assuming approximately
the same pressures in the oxygen plenum 176 and the fuel manifold
186. This will aid the combustion of the liquid hydrocarbon fuel
whenever it is injected, without local quenching by the cold oxygen
gas flowing into the oxygen heater 44.
The divert thrusters 46, 48, 50, and 52 require carefully
controlled, highly predictable mass flow rates of the liquid
hydrocarbon fuel and the oxygen gas in order to meet the tolerance
requirements for thrust and mixture ratio. The pressure and the
temperature of the hot oxygen gas must both be carefully controlled
in order to provide the desired oxygen mass flow rate to the divert
thrusters 46, 48, 50, and 52. The pressure in the oxygen heater 44
will be accurately controlled at all times by the action of the
check valves 188 and 194, which will keep the pressure in the
oxygen heater 44 approximately equal to the pressure of the oxygen
plenum 176, less the cracking pressure of the check valves 188 and
194.
However, the temperature of the oxygen gas flow from the oxygen
plenum 176 to the oxygen heater 44 will be more variable due to the
effects of two factors. The first of these factors is the ambient
temperature. The second factor is the temperature decay inside the
oxygen gas storage tanks 40 and 42 due to isentropic expansion as
the oxygen gas is used.
The base liquid hydrocarbon fuel flow rate into the oxygen heater
44 will be controlled by the check valve 188. The temperature in
the oxygen heater 44 will be continuously monitored by a
temperature sensor 196, which supplies a temperature signal which
is continuously monitored by the projectile computer control system
66. If the temperature in the oxygen heater 44 needs to be
increased, the solenoid pressure control valve 190 will be pulsed
by the projectile computer control system 66 to provide additional
liquid hydrocarbon fuel to the oxygen heater 44 via the check valve
192, where the liquid hydrocarbon fuel will immediately be
combusted in the high pressure, high temperature oxygen gas.
The mixture ratio of oxygen gas to liquid hydrocarbon fuel injected
into the oxygen heater 44 will be selected to produce a hot oxygen
gas temperature which will be sufficiently hot to ensure rapid
ignition of the divert thrusters 46, 48, 50, and 52 whenever the
liquid hydrocarbon fuel and the hot oxygen gas are mixed in the
divert thrusters 46, 48, 50, and 52. The check valves 188 and 192
on the liquid hydrocarbon fuel lines supplying the oxygen heater 44
and the check valve 194 on the oxygen gas line supplying the oxygen
heater 44 will prevent any backflow of the hot oxygen gas
therethrough into the fuel manifold 186 or the oxygen plenum 176 in
case of a brief period of excess pressure in the oxygen heater
44.
The check valves 188, 192, and 194 are sized to provide a mass flow
ratio of oxygen gas and liquid hydrocarbon fuel which, when
combusted in the oxygen heater 44, will produce a gas temperature
of approximately 1200.degree. F. to 2000.degree. F. The chemical
composition of the hot oxygen gas in the oxygen heater 44 will be
primarily oxygen, with only small amounts of products of combustion
(hence its shorthand description as "hot oxygen gas").
For the example given above in which an oxygen to fuel mass flow
ratio of 40:1 is supplied to the oxygen heater 44, the hot gas
produced by the oxygen heater 44 will consist of approximately 89
percent oxygen (in the form of O.sub.2), 5.5 percent water (H.sub.2
O), and 5.5 percent carbon dioxide (CO.sub.2). For this oxygen to
fuel mass flow ratio, the gas temperature within the oxygen heater
44 will be approximately 1850.degree. F.
It will thus be appreciated by those skilled in the art that the
gas temperature of the hot oxygen gas supplied by the oxygen heater
44, as well as its pressure, may be carefully controlled by the
projectile computer control system 66 to produce the desired
characteristics. (Note that the operation of the propulsion systems
of FIGS. 6 through 10, which will be discussed below, each have the
same arrangements and operational characteristics of the oxygen
heater 44, and are controlled in the same manner as has been
described in reference to the propulsion system of FIG. 5.)
Referring once again to FIG. 5, liquid hydrocarbon fuel is supplied
under pressure from the fuel manifold 186 to the fuel valves 140,
142, 144, and 146 of the divert thrusters 46, 48, 50, and 52,
respectively. In a similar manner, hot oxygen gas is supplied under
pressure from the oxygen heater 44 to the oxidizer valves 148, 150,
152, and 154 of the divert thrusters 46, 48, 50, and 52,
respectively. (Note that for the operation of the propulsion
systems of FIGS. 6 through 10, which will be discussed below,
liquid hydrocarbon fuel and hot oxygen gas are supplied to the
divert thrusters 46, 48, 50, and 52 in the same manner as is
accomplished in the propulsion system of FIG. 5.)
The rest of the propulsion system in FIG. 5 is quite simple in
construction. Oxygen gas is supplied under pressure from the oxygen
plenum 176 to the control valves 156, 158, 160, 162, 164, and 166
of the attitude control thrusters 54, 56, 58, 60, 62, and 64,
respectively. The attitude control thrusters 54, 56, 58, 60, 62,
and 64 are jets in operation, since the release of the oxygen gas
supplied from the oxygen plenum 176 is their sole source of
thrust.
This completes the description of the propulsion system of FIG. 5,
which may be characterized as having its liquid hydrocarbon fuel
storage tanks 36 and 38 pressurized by oxygen gas from the oxygen
plenum 176, and as having its attitude control thrusters 54, 56,
58, 60, 62, and 64 operated by (cold) oxygen gas from the oxygen
plenum 176.
With regard generally to FIGS. 6 through 10, a number of the
components of these propulsion systems and their methods of
operation are identical to the construction and the method of
operation of the propulsion system of FIG. 5, as made clear above
with reference to FIG. 5. Thus, only those portions of the
propulsion systems of FIGS. 6 through 10 which differ in
construction and operation from the propulsion system of FIG. 5
will be discussed below in detail.
Referring now specifically to FIG. 6, hot oxygen gas is supplied
under pressure from the oxygen heater 44 to the control valves 156,
158, 160, 162, 164, and 166 of the attitude control thrusters 54,
56, 58, 60, 62, and 64, respectively. The attitude control
thrusters 54, 56, 58, 60, 62, and 64 are again jets in operation,
since the release of the hot oxygen gas supplied from the oxygen
heater 44 is their sole source of thrust.
This completes the description of the propulsion system of FIG. 6,
which may be characterized as having its liquid hydrocarbon fuel
storage tanks 36 and 38 pressurized by oxygen gas from the oxygen
plenum 176, and as having its attitude control thrusters 54, 56,
58, 60, 62, and 64 operated by hot oxygen gas from the oxygen
heater 44.
Referring now specifically to FIG. 7, instead of the attitude
control thrusters 54, 56, 58, 60, 62, and 64, which operate
functionally as jets, the propulsion system depicted instead uses
true bipropellant thrusters similar to the divert thrusters 46, 48,
50, and 52. Six attitude control thrusters 200, 202, 204, 206, 208,
and 210 are thus used.
The attitude control thrusters 200, 202, 204, 206, 208, and 210
each have a fuel valve and an oxidizer valve. Specifically, the
attitude control thruster 200 has a fuel valve 212 and an oxidizer
valve 224, the attitude control thruster 202 has a fuel valve 214
and an oxidizer valve 226, the attitude control thruster 204 has a
fuel valve 216 and an oxidizer valve 228, the attitude control
thruster 206 has a fuel valve 218 and an oxidizer valve 230, the
attitude control thruster 208 has a fuel valve 220 and an oxidizer
valve 232, and the attitude control thruster 210 has a fuel valve
222 and an oxidizer valve 234. Operation of the fuel valves 212,
214, 216, 218, 220, and 222 and the oxidizer valves 224, 226, 228,
230, 232, and 234 are all controlled by the projectile computer
control system 66.
Liquid hydrocarbon fuel is supplied under pressure from the fuel
manifold 186 to the fuel valves 212, 214, 216, 218, 220, and 222 of
the attitude control thrusters 200, 202, 204, 206, 208, and 210,
respectively. In a similar manner, hot oxygen gas is supplied under
pressure from the oxygen heater 44 to the oxidizer valves 224, 226,
228, 230, 232, and 234 of the attitude control thrusters 200, 202,
204, 206, 208, and 210, respectively.
This completes the description of the propulsion system of FIG. 7,
which may be characterized as having its liquid hydrocarbon fuel
storage tanks 36 and 38 pressurized by oxygen gas from the oxygen
plenum 176, and as having its attitude control thrusters 200, 202,
204, 206, 208, and 210 being bipropellant operated using liquid
hydrocarbon fuel from the fuel manifold 186 and hot oxygen gas from
the oxygen heater 44.
Referring now specifically to FIG. 8, the inlets of the liquid
hydrocarbon fuel storage tanks 36 and 38 are not pressurized by
oxygen gas from the oxygen plenum 176, but rather by an independent
helium gas system first illustrated in FIG. 8. Helium gas is stored
in a helium gas storage tank 236 at high pressure, typically on the
order of approximately 10,000 psia. Pressure in the helium gas
storage tank 236 is monitored by a pressure transducer 238, which
supplies a pressure signal which is monitored by the projectile
computer control system 66.
For safety purposes, a burst disk 240 is installed on the outlet
from the helium gas storage tank 236. The burst disk 240 is
designed to rupture at a pressure higher than the maximum helium
gas pressure in storage or operation of the propulsion system, but
lower than the burst pressure of the helium gas storage tank 236.
In the event of the propulsion system being engulfed in a fire, the
pressure in the helium gas storage tank 236 will rise until the
burst disk 240 ruptures. The helium gas will then be discharged
from the helium gas storage tank 236, thereby avoiding the rupture
and fragmentation of the helium gas storage tank 236.
A pyrovalve 242 is installed between the outlet from the helium gas
storage tank 236 and a solenoid pressure control valve 244. The
pyrovalve 242 is made for one-time operation, and once opened by a
signal from the projectile computer control system 66 will remain
open to allow the passage of helium gas therethrough. The pyrovalve
242 functions to isolate the liquid hydrocarbon fuel storage tanks
36 and 38 from the helium gas during storage.
Once the pyrovalve 242 has been opened, the solenoid pressure
control valve 244 will be opened by the projectile computer control
system 66 to allow helium gas to flow into the inlets of the liquid
hydrocarbon fuel storage tanks 36 and 38. As an added safety
feature, note that if the pyrovalve 242 were fired accidentally, or
by the occurrence of a fire, high pressure helium gas would still
be isolated from the liquid hydrocarbon fuel storage tanks 36 and
38 by the solenoid pressure control valve 244. Pressure in the
inlets of the liquid hydrocarbon fuel storage tanks 36 and 38 is
monitored by a pressure transducer 246, which supplies a pressure
signal which is monitored by the projectile computer control system
66. Using this pressure signal, the projectile computer control
system 66 will operate the solenoid pressure control valve 244 to
control the pressure of helium gas in the inlets of the liquid
hydrocarbon fuel storage tanks 36 and 38.
Typically, pressure in the inlets of the liquid hydrocarbon fuel
storage tanks 36 and 38 is maintained at approximately the same
pressure as the pressure in the oxygen plenum 176 (typically 1000
psi). Whenever the pressure in the inlets of the liquid hydrocarbon
fuel storage tanks 36 and 38 drops below the desired pressure, the
projectile computer control system 66 will cause the solenoid
pressure control valve 244 to be actuated for a short duration,
typically a few milliseconds, to allow more helium gas to be
injected into the inlets of the liquid hydrocarbon fuel storage
tanks 36 and 38, thereby raising the pressure in the liquid
hydrocarbon fuel storage tanks 36 and 38.
(Note that the propulsion systems of FIGS. 9 and 10, which will be
discussed below, each have the same helium gas pressurization of
the liquid hydrocarbon fuel storage tanks 36 and 38, as well as the
same arrangements of the pressure transducer 238, the burst disk
240, the pyrovalve 242, the solenoid pressure control valve 244,
and the pressure transducer 246 as contained in the propulsion
system of FIG. 8, and are operated in the same manner by their
relative projectile computer control systems 66.)
The rest of the propulsion system of FIG. 8 may now be briefly
described. Rather than use only the oxygen plenum 176 as in FIG. 5,
the propulsion system of FIG. 8 also uses an attitude control
system plenum 248 to distribute oxygen gas supplied from the oxygen
gas storage tanks 40 and 42. When the pyrovalve 172 has been opened
by the projectile computer control system 66, oxygen gas will also
be supplied to a solenoid pressure control valve 250.
The solenoid pressure control valve 250 will be opened by the
projectile computer control system 66 to allow oxygen gas to flow
into the attitude control system plenum 248. Pressure in the
attitude control system plenum 248 is monitored by a pressure
transducer 252, which supplies a pressure signal which is monitored
by the projectile computer control system 66. Using this pressure
signal, the projectile computer control system 66 will operate the
solenoid pressure control valve 250 to control the pressure of
oxygen gas in the attitude control system plenum 248.
Typically, pressure in the attitude control system plenum 248 is
also maintained at approximately 1000 psi. Whenever the pressure in
the attitude control system plenum 248 drops below the desired
pressure, the projectile computer control system 66 will cause the
solenoid pressure control valve 250 to be actuated for a short
duration, typically a few milliseconds, to allow more oxygen gas to
be injected into the attitude control system plenum 248, thereby
raising the pressure in the attitude control system plenum 248.
Oxygen gas is supplied under pressure from the attitude control
system plenum 248 to the control valves 156, 158, 160, 162, 164,
and 166 of the attitude control thrusters 54, 56, 58, 60, 62, and
64, respectively. The attitude control thrusters 54, 56, 58, 60,
62, and 64 are jets in operation, since the release of the oxygen
gas supplied from the attitude control system plenum 248 is their
sole source of thrust. Unlike the propulsion system of FIG. 5, the
propulsion system of FIG. 8 uses the attitude control system plenum
248 in addition to the oxygen plenum 176, and thus includes the
additional components necessary to the operation of the attitude
control system plenum 248.
This completes the description of the propulsion system of FIG. 8,
which may be characterized as having its liquid hydrocarbon fuel
storage tanks 36 and 38 pressurized by helium gas, and as having
its attitude control 54, 56, 58, 60, 62, and 64 operated by (cold)
oxygen gas from the attitude control system plenum 248.
Referring now specifically to FIG. 9, oxygen gas is supplied under
pressure from the oxygen plenum 176 to the control valves 156, 158,
160, 162, 164, and 166 of the attitude control thrusters 54, 56,
58, 60, 62, and 64, respectively. The attitude control thrusters
54, 56, 58, 60, 62, and 64 are jets in operation, since the release
of the oxygen gas supplied from the oxygen plenum 176 is their sole
source of thrust.
This completes the description of the propulsion system of FIG. 9,
which may be characterized as having its liquid hydrocarbon fuel
storage tanks 36 and 38 pressurized by helium gas, and as having
its attitude control 54, 56, 58, 60, 62, and 64 operated by (cold)
oxygen gas from the oxygen plenum 176. Unlike the propulsion system
of FIG. 8, the propulsion system of FIG. 9 does not use the
attitude control system plenum 248 in addition to the oxygen plenum
176, and therefore is simplified in comparison to the propulsion
system of FIG. 8.
Referring now specifically to FIG. 10, hot oxygen gas is supplied
under pressure from the oxygen heater 44 to the control valves 156,
158, 160, 162, 164, and 166 of the attitude control thrusters 54,
56, 58, 60, 62, and 64, respectively. The attitude control
thrusters 54, 56, 58, 60, 62, and 64 are again jets in operation,
since the release of the hot oxygen gas supplied from the oxygen
heater 44 is their sole source of thrust.
This completes the description of the propulsion system of FIG. 10,
which may be characterized as having its liquid hydrocarbon fuel
storage tanks 36 and 38 pressurized by helium gas, and as having
its attitude control 54, 56, 58, 60, 62, and 64 operated by hot
oxygen gas from the oxygen heater 44.
Although only six different embodiments of the propulsion system of
the present invention have been illustrated herein, it will at once
be appreciated by those skilled in the art that the various
different design features of the present invention illustrated in
those six embodiments may be combined in a wide variety of
different permutations. Such different combinations will be readily
apparent from the exemplary illustrations presented in the above
description of the propulsion system of the present invention.
It may therefore be appreciated from the above detailed description
of the preferred embodiment of the present invention that it
teaches an improved propulsion system useable for divert and
attitude control propulsion systems, and a related method for use
thereof, which system and method exclusively use propellants which
are non-toxic. As such, both the fuel and the oxidizer of the
present invention are non-toxic, thereby eliminating one of the
most serious drawbacks of previously known liquid or gel
bipropellants or monopropellants. The propellants used by the
propulsion system and the related method of the present invention
are not subject to detonation in a fire, when hit by a stray
bullet, or when subjected to high temperature.
The propulsion system and the related method of the present
invention are efficiently operable in an on-off pulsatile manner
unlike solid propellant systems. The improved propulsion system of
the present invention also presents excellent center-of-gravity
control, such that the characteristics of the missile in which the
propulsion system of the present invention is installed will also
feature excellent dynamic center-of-gravity characteristics. The
weight of the improved propulsion system of the present invention
is also relatively light in comparison to solid propulsion systems,
and in fact is comparable to or less than the weight of previously
known liquid bipropellant or monopropellant systems.
The propulsion system of the present invention is also of a
construction which is both durable during operation, and long
lasting in a storage situation, and which requires little or no
maintenance to be provided throughout the time that it is stored.
The propulsion system of the present invention is of comparable
cost to previously known liquid bipropellant and monopropellant
propulsion systems, and to solid propellant propulsion systems,
thereby affording the propulsion system of the present invention
the broadest possible market. Finally, all of the aforesaid
advantages and objectives of the propellant system and the related
method of use of the present invention are achieved without
incurring any substantial relative disadvantage.
Although an exemplary embodiment of the present invention has been
shown and described with reference to particular embodiments and
applications thereof, it will be apparent to those having ordinary
skill in the art that a number of changes, modifications, or
alterations to the invention as described herein may be made, none
of which depart from the spirit or scope of the present invention.
All such changes, modifications, and alterations should therefore
be seen as being within the scope of the present invention.
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